1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine airfoil with shaped film cooling holes.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
Airfoils in stator vanes and rotor blades make use of film cooling holes to discharge a layer of film cooling air onto the external surface of the airfoil to form an insulation layer of cooler air against the hot gas stream passing over the airfoil surface and to keep the airfoil surface at or below a certain allowable metal temperature. Earlier film cooling holes were straight and at a constant diameter. More recent film cooling holes are shaped with a diffusion opening that produces greater film coverage. Straight film cooling holes pass straight though the airfoil wall at a constant diameter and exit at an angle to the airfoil surface. FIGS. 1-7 show prior art straight film holes. Some of the cooling air is ejected directly into the mainstream gas flow and causes turbulence, coolant dilution and a loss of downstream film effectiveness. Straight film holes form an opening on the airfoil surface (referred to as the hole breakout) in al elliptical shape that will induce a high stress issue on the airfoil surface.
FIG. 8 shows a prior art shaped film cooling hole with a diffusion section formed downstream from an inlet metering section. The FIG. 8 shaped film hole is referred to as a 10×10×10 film hole because the diffusion section has a 10 degree slope on each of the two side walls and the one downstream wall. The upstream wall does not slope and thus provides no diffusion to the film hole. During engine operation, hot gas is frequently captured within an upper corner of the film hole opening (#11 in FIG. 1) and causes shear mixing with the cooling air flowing through from the film hole. This causes a reduction of film cooling effectiveness for the film cooling hole. Also, internal flow separation occurs within the diffusion section of the film hole (#12 in FIG. 10) at a junction between the constant diameter metering section and the diffusion section.